The present invention relates generally to gas turbine engines, and, more specifically, to turbine blade damping.
A gas turbine engine includes a turbine rotor or disk in which a plurality of circumferentially spaced apart turbine blades are supported around the perimeter. Each blade includes a hollow airfoil over which combustion gases flow during operation, with a platform being disposed at the root of the airfoil to define an inner boundary for the combustion gases. Extending radially below the platform is an integral shank and a corresponding dovetail therebelow. The dovetail may be configured as an axial-entry or a circumferential-entry dovetail, with the former being mounted in a complementary dovetail slot extending axially through the perimeter of the rotor disk.
During operation, the rotor disk is rotated by the extraction of energy from the hot combustion gases at the airfoils, and is therefore subject to vibration caused by rotation of the blades and aerodynamic loading of the airfoils. Blade vibration can occur at multiple natural frequencies, and corresponding modes, as excited by the speed of rotation and aerodynamic stimuli. Since a turbine operates over a range of rotary speed, different modes of vibration may be excited differently, and are therefore subject to different amounts of vibratory amplitude.
Accordingly, turbine rotor blades are specifically designed to minimize vibratory motion during operation while achieving a correspondingly long useful life. The high cycle fatigue strength of a turbine blade is one contributor to blade life, and is compromised when fatigue cracks appear near the end of blade life. High cycle fatigue cracks are initiated over the cumulative effect of vibratory motion of the blade during operation and typically occur in high stress regions of the blade, such as the airfoil, dovetail, or shank.
In order to improve the high cycle fatigue life of a turbine blade, vibration dampers are provided below the blade platforms to frictionally dissipate vibratory energy and reduce the corresponding amplitude of vibration during operation. A typical vibration damper is a thin sheet metal component having a trapezoidal profile which is loosely retained or trapped under adjoining platforms to bridge the axial splitline therebetween.
The damper is trapped radially between the adjacent platforms in corresponding pairs of lugs extending circumferentially outwardly from the opposing blade shanks. Under centrifugal force, the damper radially engages the underside of the blade platforms and conforms thereto for providing a frictional interface therebetween and a fluid seal at the splitline. The dampers are sized for achieving sufficient mass for effectively dissipating vibratory energy of the blades carried through the blade platforms.
However, the thin dampers must also be retained axially under the platforms to prevent undesirable liberation therefrom. An improved turbine blade vibration damper in the shape of an hourglass includes symmetrical, concave side notches extending longitudinally between a pair of opposite end tabs in a unitary sheet metal component. The symmetrical configuration of the damper permits its correct assembly between adjacent platforms in any one of the four possible installation orientations. When installed, one of the two side notches conforms with a corresponding convex bulge from the blade shank below the convex, suction side of the airfoil through which a cooling air passage extends radially from the airfoil and through the shank and dovetail for receiving cooling air during operation.
However, testing of this improved design has shown that under certain circumstances the thin damper may slide axially sufficiently to disengage the side notch from the shank bulge causing undesirable distortion of the damper, which in turn may lead to damage or liberation thereof.
Accordingly, it is desired to provide an improved turbine blade vibration damper having sufficient damping mass with self retention for preventing damage and liberation thereof during operation.